Shaft bearing positioning in a gas turbine engine

ABSTRACT

A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and a fan including a plurality of fan blades located upstream of the engine core. The fan has a fan diameter in the range from 240 cm to 280 cm. The turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The turbine includes a total of three sets of turbine blades. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings located downstream of a leading edge of a lowest pressure turbine blade of the turbine at a root of the blade.

This is a Continuation of application Ser. No. 16/809,984 filed Mar. 5,2020, which in turn claims the benefit of GB 1918781.4, filed Dec. 19,2019. The disclosure of the prior applications is hereby incorporated byreference herein in its entirety.

The present disclosure relates to the mounting of a core shaft within agas turbine engine for an aircraft, and in particular to bearingpositioning and how such a shaft may be arranged and mounted so as tomanage vibrational and resonance effects.

As used herein, a range “of value X to value Y”, “from value X to valueY” or “between value X and value Y”, or the likes, denotes an inclusiverange; including the bounding values of X and Y. As used herein, theterm “axial plane” denotes a plane extending along the length of anengine, parallel to and containing an axial centreline of the engine,and the term “radial plane” denotes a plane extending perpendicular tothe axial centreline of the engine, so including all radial lines at theaxial position of the radial plane. Axial planes may also be referred toas longitudinal planes, as they extend along the length of the engine. Aradial distance or an axial distance is therefore a distance in a radialor axial plane, respectively.

According to a first aspect, there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the compressor is the lowest pressure compressor of theengine and the turbine is the lowest pressure turbine of the engine andhas a lowest pressure set of blades each blade of the lowest pressureset of blades having a mass, m, a radius at blade mid-height, r, and anangular velocity at cruise, co. The engine further comprises a fanlocated upstream of the engine core, the fan comprising a plurality offan blades; and a gearbox that receives an input from the core shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft.

The engine core further comprises three bearings arranged to support thecore shaft, the three bearings comprising a forward bearing and tworearward bearings, with a minor span (S) being defined as the axialdistance between the two rearward bearings. A first blade to bearingratio of:

$\frac{{the}{minor}{span}(S)}{mr{\omega^{2}\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$has a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s².

The inventor appreciated that the engine should not be linearly scaledup with a force increase (mrω² providing a measure of force), but ratherthat the minor span length (S) should be increased as little as possibleso as to relatively reduce engine length and weight, so allowing theefficiency gains to be increased by avoiding the additional weight, andto avoid the development of unwanted whirl modes within the minor span.Whilst conventional wisdom suggests that a larger minor span isdesirable to improve reaction of forces from the low pressure turbine,the inventor found that the risk of introducing whirl modes, and theintroduction of more length and weight, counterbalanced the forcereaction benefits and that the first blade to bearing ratio shouldtherefore be maintained within the specified range.

The first blade to bearing ratio may be in the range from 3.0×10⁻⁶ to7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 4.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 3.0×10⁻⁶ to 4.5×10⁻⁶kg⁻¹·rad⁻²·s², and further optionally from 4.5×10⁻⁶ to 6.5×10⁻⁶kg⁻¹·rad⁻²·s².

A second blade to bearing ratio of:

$\frac{{the}{minor}{span}(S)}{m \times {r\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$may have a value in the range from 0.8 to 6.0 kg⁻¹, optionally from 0.8to 5.0 kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0kg⁻¹, optionally from 0.8 to 2.0 kg⁻¹.

According to an second aspect, there is provided a gas turbine enginefor an aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine andhas a lowest pressure set of blades, each blade of the lowest pressureset of blades having a mass, m, a radius at blade mid-height, r, and thecompressor is the lowest pressure compressor of the engine. The enginefurther comprises a fan located upstream of the engine core, the fancomprising a plurality of fan blades; and a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support thecore shaft, the three bearings comprising a forward bearing and tworearward bearings, with a minor span (S) being defined as the axialdistance between the two rearward bearings. A second blade to bearingratio of:

$\frac{{the}{minor}{span}(S)}{m \times {r\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$has a value in the range from 0.8 to 6.0 kg⁻¹.

The second blade to bearing ratio may be in the range from 0.8 to 5.0kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0 kg⁻¹,optionally from 0.8 to 2.0 kg⁻¹.

Each blade of the lowest pressure set of blades has an angular velocityat cruise, co, and a first blade to bearing ratio of:

$\frac{{the}{minor}{span}(S)}{mr{\omega^{2}\left( {{{for}a{blade}{of}{the}{lowest}{pressure}{set}19c},{19d}} \right)}}$

may have a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from4.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and furtheroptionally from 4.5×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

In the first or second aspects, one or more of the following featuresmay be present:

The minor span, S, may be in the range from 250 mm to 350 mm. In variousembodiments, the minor span may be greater than or equal to any of 250mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor spanmay be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335mm.

The length of the core shaft (L) may be in the range from 1800 mm to2900 mm, optionally from 2000 mm to 2900 mm, further optionally from2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

The value of blade mass, m, multiplied by blade radius at mid-height, r,may be in the range from 180 to 280 kg·mm.

The gearbox may have a gear ratio greater than 3, and optionally in therange from 3.1 to 3.8.

The core shaft may have a running speed in the range from 1500 rpm to6200 rpm.

The running speed of the core shaft at cruise may be in the range of5400-5700 rpm, and optionally of 5500-5600 rpm, at cruise.

The running speed of the core shaft under maximum take-off (MTO)conditions may be in the range of 5800-6200 rpm, and optionally of5900-6100 rpm.

A length ratio (S/L) of the minor span between the two rearward bearingsto the core shaft length may be equal to or less than 0.14, or equal toor less than 0.13, or equal to or less than 0.12. The length ratio S/Lmay be equal to or greater than 0.05, or equal to or greater than 0.06,or equal to or greater than 0.07, or equal to or greater than 0.08. Forexample, the length ratio S/L may be in the range from 0.05 to 0.14,optionally in the range from 0.05 to 0.13, optionally in the range from0.06 to 0.13, and optionally in the range from 0.08 to 0.13.

The mass, m, of a blade of the lowest pressure set of blades may be inthe range from 0.2 to 0.6 kg.

The radius, r, of a blade of the lowest pressure set of blades may be inthe range from 400 to 600 mm.

Each blade of the lowest pressure set of blades may have an angularvelocity at cruise, co in the range from 560 to 600 rad·s⁻¹.

The length of the core shaft may be in the range from 1800 mm to 2900 mmor 2750 mm. The minor span may be in the range from 250 mm to 350 mm,and optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The turbine has a turbine length defined between the leading edge of itsmost upstream blades and a trailing edge of its most downstream blades.A minor span to turbine length ratio of:

$\frac{{minor}{span}}{{turbine}{length}}$may be equal to or less than 1.05, optionally equal to or less than1.00, optionally equal to or less than 0.95. The minor span to turbinelength ratio may be equal to or greater than 0.70, or equal to orgreater than 0.75, or equal to or greater than 0.80, or equal to orgreater than 0.85. For example the minor span to turbine length ratiomay be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00,optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionallyfrom 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to0.95.

The rearward bearings may be positioned axially level with or rearwardof a leading edge of a lowest pressure turbine blade of the turbine atthe root of the blade.

The rearward bearings may be positioned axially level with or rearwardof a trailing edge of a turbine blade of a third set of turbine bladesfrom the front of the turbine, at the root of the blade. In suchembodiments, the turbine may comprise four sets of turbine blades, andoptionally may have a total of four sets of turbine blades.

The forwardmost bearing of the rearward bearings may have a bearingstiffness in the range of 30 kN/mm to 100 kN/mm.

A stiffness ratio of the bearing stiffness at the forwardmost rearwardbearing to the distance between the two rearward bearings may be in therange from 0.08 to 0.5 kN/mm², optionally in the range from 0.08 to 0.40kN/mm², optionally in the range from 0.08 to 0.30 kN/mm², optionally inthe range from 0.08 to 0.20 kN/mm², optionally in the range from 0.09 to0.40 kN/mm², optionally in the range from 0.15 to 0.50 kN/mm²,optionally in the range from 0.15 to 0.40 kN/mm², and further optionallyin the range from 0.15 to 0.30 kN/mm².

According to a third aspect, there is provided a method of operation ofa gas turbine engine for an aircraft, the engine comprising an enginecore comprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor, and wherein the compressor is the lowestpressure compressor of the engine and the turbine is the lowest pressureturbine of the engine and has a lowest pressure set of blades, eachblade of the lowest pressure set of blades having a mass, m, a radius atblade mid-height, r, and an angular velocity at cruise, co. The enginecore further comprises three bearings arranged to support the coreshaft, the three bearings comprising a forward bearing and two rearwardbearings, with a minor span (S) being defined as the axial distancebetween the two rearward bearings. The engine further comprises a fanlocated upstream of the engine core, the fan comprising a plurality offan blades; and a gearbox arranged to receive an input from the coreshaft and to output drive to the fan so as to drive the fan at a lowerrotational speed than the core shaft.

The method comprises operating the engine such that a first blade tobearing ratio of:

$\frac{{the}{minor}{span}(S)}{mr{\omega^{2}\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$has a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s² atcruise.

In various embodiments the method may comprise operating the engine suchthat the first blade to bearing ratio may be in the range from 3.0×10⁻⁶to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 4.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and further optionally from 4.5×10⁻⁶to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

The method may further comprise operating the core shaft at a runningspeed in the range from 1500 rpm to 6200 rpm, and optionally wherein therunning speed of the core shaft may be one or more of:

-   -   (i) in the range of 5400-5700 rpm, and optionally of 5500-5600        rpm, at cruise; and/or    -   (ii) in the range of 5800-6200 rpm, and optionally of 5900-6100        rpm, under maximum take-off conditions.

The engine used to implement the method of the third aspect may be asdescribed in the first and/or second aspect.

According to a fourth aspect, there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine,the core shaft has a running speed in the range from 1500 rpm to 6200rpm, and the compressor is the lowest pressure compressor of the engine.The engine further comprises a fan located upstream of the engine core,the fan comprising a plurality of fan blades; and a gearbox thatreceives an input from the core shaft and outputs drive to the fan so asto drive the fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support thecore shaft, the three bearings comprising a forward bearing and tworearward bearings, the core shaft having a length (L) between theforward bearing and the rearmost rearward bearing in the range from 1800mm to 2900 mm or 2750 mm, and a minor span (S) between the two rearwardbearings in the range from 250 mm to 350 mm.

As a result, there may be no primary resonance of the core shaft betweenthe forward bearing and the forwardmost rearward bearings within therunning speed range of the core shaft.

The length (L) and minor span (S) may be selected as appropriate for thedesired running speed range to avoid such primary resonances.

In various embodiments the length of the core shaft (L) may be in therange from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The inventor appreciated that the increased length of a core shaft whena gas turbine engine is scaled up may lead to a resonance frequency ofthe core shaft lying in or near the engine running range. Simply scalingup a known engine may therefore lead to increased risks ofresonance-induced damage—the longer core shaft may be problematic. Theinventor discovered that selecting the length, L, and minor span, S, asappropriate for a given running speed range may facilitate avoidance ofdeleterious whirl modes, and may reduce damage to the engine in use.

The lower bound of 1500 rpm on the core shaft running speed may be theminimum running speed under ground idle conditions and/or the upperbound of 6200 rpm on the core shaft running speed may be the upper boundon maximum take-off running speed.

According to a fifth aspect, there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine,and the compressor is the lowest pressure compressor of the engine. Theengine further comprises a fan located upstream of the engine core, thefan comprising a plurality of fan blades and having a fan diameter inthe range from 330 cm to 380 cm; and a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft, the gearbox having agear ratio in the range from 3.1 to 3.8.

The engine core further comprises three bearings arranged to support thecore shaft, the three bearings comprising a forward bearing and tworearward bearings, the core shaft having a length between the forwardbearing and the rearmost rearward bearing in the range from 1800 to 2900mm or 2750 mm, and a minor span between the two rearward bearings in therange from 250 mm to 350 mm, such that there is no primary resonance ofthe core shaft within a running speed range of the core shaft.

The skilled person would appreciate that the fifth aspect shares aninventive concept with the fourth aspect, as the combination of fandiameter and gear ratio is linked to core shaft running speed. For aparticular design of aircraft, a given combination of fan diameter andgear ratio may be used to infer an intended core shaft running speedrange.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The core shaft running speed range may be from 1500 rpm to 6200 rpm.

In the fourth or fifth aspects, one or more of the following featuresmay be present:

The turbine has a turbine length defined between the leading edge of itsmost upstream blades and a trailing edge of its most downstream blades.A minor span to turbine length ratio of:

$\frac{{minor}{span}}{{turbine}{length}}$may be equal to or less than 1.05. The minor span to turbine lengthratio may be equal to or less than 1.00, optionally equal to or lessthan 0.95. The minor span to turbine length ratio may be equal to orgreater than 0.70, or equal to or greater than 0.75, or equal to orgreater than 0.80, or equal to or greater than 0.85. For example theminor span to turbine length ratio may be in the range from 0.70 to1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95,optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionallyfrom 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to1.00, and further optionally from 0.85 to 0.95 in the range from 0.85 to0.95.

The turbine may comprise four sets of turbine blades. The two rearwardbearings may both be located downstream of the trailing edge of aturbine blade of the third set of turbine blades from the front of theturbine, at the root of the blade. The turbine may comprise a total offour sets of turbine blades.

The two rearward bearings may be located downstream of the leading edgeof the lowest pressure (most downstream) turbine blade of the turbine atthe root of the blade.

The length of the core shaft (L) may be in the range from 1800 mm to2900 mm, optionally from 2000 mm to 2900 mm, further optionally from2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

A length ratio (S/L) of the minor span between the two rearward bearingsto the core shaft length may be equal to or less than 0.14.

The length ratio may be equal to or less than 0.13, or equal to or lessthan 0.12. The length ratio may be equal to or greater than 0.05, orequal to or greater than 0.06, or equal to or greater than 0.07, orequal to or greater than 0.08. For example, the length ratio may be inthe range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13,optionally in the range from 0.06 to 0.13, and further optionally in therange from 0.08 to 0.13.

The forwardmost bearing of the rearward bearings may have a bearingstiffness in the range of 30 kN/mm to 100 kN/mm.

A stiffness ratio of the bearing stiffness at the forwardmost rearwardbearing to the distance between the two rearward bearings may be in therange from 0.08 to 0.5 kN/mm², optionally in the range from 0.08 to 0.40kN/mm², optionally in the range from 0.08 to 0.30 kN/mm², optionally inthe range from 0.08 to 0.20 kN/mm², optionally in the range from 0.09 to0.40 kN/mm², optionally in the range from 0.15 to 0.50 kN/mm²,optionally in the range from 0.15 to 0.40 kN/mm², and further optionallyin the range from 0.15 to 0.30 kN/mm².

The lowest pressure turbine of the engine has a lowest pressure set ofblades, each blade of the lowest pressure set of blades having a mass,m, a radius at blade mid-height, r, and an angular velocity at cruise,ω.

A first blade to bearing ratio of:

$\frac{{the}{minor}{span}(S)}{mr{\omega^{2}\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$may have a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from4.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and furtheroptionally from 4.5×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

A second blade to bearing ratio of:

$\frac{{the}{minor}{span}(S)}{m \times {r\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$may have a value in the range from 0.8 to 6.0 kg⁻¹, optionally from 0.8to 5.0 kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0kg⁻¹, optionally from 0.8 to 2.0 kg⁻¹.

According to a sixth aspect, there is provided a method of operation ofa gas turbine engine for an aircraft. The gas turbine engine comprisesan engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor, and wherein the turbine is thelowest pressure turbine of the engine. The engine core further comprisesthree bearings arranged to support the core shaft, the three bearingscomprising a forward bearing and two rearward bearings, the core shafthaving a length (L) between the forward bearing and the rearmostrearward bearing in the range from 1800 mm to 2900 mm, and a minor span(S) between the two rearward bearings in the range from 250 mm to 350mm. The engine further comprises a fan located upstream of the enginecore, the fan comprising a plurality of fan blades; and a gearboxarranged to receive an input from the core shaft and to output drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft

The method comprises operating the engine such that the core shaft has arunning speed in the range from 1500 rpm to 6200 rpm, and wherein thereis no primary resonance of the core shaft within the running speed rangeof the core shaft.

The length of the core shaft (L) may be in the range from 1800 mm to2750 mm, or from 2000 to 2750, or from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The engine used to implement the method may be as described in thefourth and/or fifth aspects.

According to a seventh aspect, there is provided a method of designing agas turbine engine for an aircraft. The engine comprises an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor, and wherein the turbine is the lowestpressure turbine of the engine, the core shaft has a running speed inthe range from 1500 rpm to 6200 rpm, and the compressor is the lowestpressure compressor of the engine. The engine further comprises a fanlocated upstream of the engine core, the fan comprising a plurality offan blades; and a gearbox that receives an input from the core shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft. The engine core further comprises threebearings arranged to support the core shaft, the three bearingscomprising a forward bearing and two rearward bearings, the core shafthaving a length (L) between the forward bearing and the rearmostrearward bearing in the range from 1800 mm to 2900 mm.

The method comprises:

-   -   selecting positions for the forward bearing and the forwardmost        bearing of the rearward bearings; and    -   lengthening the core shaft rearward of the forwardmost bearing        of the rearward bearings such that a minor span (S) defined        between the two rearward bearings is in the range from 250 mm to        350 mm, and there is no primary resonance of the core shaft        between the forward bearing and the forwardmost rearward bearing        within the running speed range of the core shaft.

The length of the core shaft (L) may be in the range from 1800 mm to2750 mm, or from 2000 to 2750, or from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The engine designed by the method of the seventh aspect may be theengine of the fourth and/or fifth aspect.

According to an eighth aspect, there is provided a gas turbine enginefor an aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine andthe compressor is the lowest pressure compressor of the engine. Theengine further comprises a fan located upstream of the engine core, thefan comprising a plurality of fan blades; and a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The engine corefurther comprises three bearings arranged to support the core shaft, thethree bearings comprising a forward bearing and two rearward bearings,and wherein the forwardmost rearward bearing has a bearing stiffness inthe range of 30 kN/mm to 100 kN/mm, the bearing stiffness being definedby the radial displacement caused by the application of a radial forceat the axial centrepoint of the bearing.

The inventor appreciated that controlling the bearing stiffness to liewithin the specified range may allow or facilitate management ofvibrational modes, so potentially reducing damage to the engine in usecaused by whirl mode displacements of the core shaft.

The core shaft has a length (L) between the forward bearing and therearmost rearward bearing and a minor span (S) between the rearwardbearings. The bearings may be arranged such that a length ratio (S/L) ofthe minor span to the core shaft length is equal to or less than 0.14,or equal to or less than 0.13, or equal to or less than 0.12. The lengthratio may be equal to or greater than 0.05, or equal to or greater than0.06, or equal to or greater than 0.07, or equal to or greater than0.08. For example, the length ratio may be in the range from 0.05 to0.14, optionally in the range from 0.05 to 0.13, optionally in the rangefrom 0.06 to 0.13, and further optionally in the range from 0.08 to0.13.

The core shaft may have a running speed range with a lower bound of 1500rpm and an upper bound of 6200 rpm.

The bearing stiffness of the forwardmost rearward bearing may be equalto or around 50 kN/mm.

The gas turbine engine may further comprise a stationary supportingstructure and a first bearing support structure. The forwardmostrearward bearing may be mounted to the stationary supporting structureby the first bearing support structure. The first bearing supportstructure may be attached to the stationary supporting structure at afirst position located axially rearward of the forwardmost rearwardbearing.

In such embodiments, the first bearing support structure may comprise aplurality of connecting members, which may be spaced circumferentiallyaround the engine axis, connecting the bearing to the stationary supportstructure. The first bearing support structure may be described as aspring bar-type support structure.

The first bearing support structure may comprise an outer race of theforwardmost rearward bearing.

The gas turbine engine of embodiments with a first bearing supportstructure may further comprise a second bearing support structure. Thesecond bearing support structure may be mounted to the stationarysupporting structure, optionally at a second position located forwardof, and at a larger radial distance from the engine axis than, the firstposition. The second bearing support structure may be connected to thefirst bearing support structure by a squeeze film damper in the regionof the forwardmost rearward bearing.

A stiffness ratio of the bearing stiffness at the forwardmost rearwardbearing to the minor span may be in the range from 0.08 to 0.5 kN/mm²,optionally in the range from 0.08 to 0.40 kN/mm², optionally in therange from 0.08 to 0.30 kN/mm², optionally in the range from 0.08 to0.20 kN/mm², optionally in the range from 0.09 to 0.40 kN/mm²,optionally in the range from 0.15 to 0.50 kN/mm², optionally in therange from 0.15 to 0.40 kN/mm², and further optionally in the range from0.15 to 0.30 kN/mm².

The length of the core shaft (L) may be in the range from 1800 mm to2900 mm, optionally from 2000 mm to 2900 mm, further optionally from2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The length of the core shaft may be in the range from 1800 mm to 2900 mmor 2750 mm. In such embodiments, the minor span may be in the range from250 mm to 350 mm, optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The length of the core shaft may be in the range from 1800 mm to 2900 mmor 2750 mm. The minor span may be in the range from 250 mm to 350 mm.The running speed of the core shaft may be in the range from 1500 rpm to6200 rpm; and/or a diameter of the fan may be in the range from 330 cmto 380 cm and the gear ratio of the gearbox is in the range from 3.1 to3.8. The length, minor span and/or running speed may be selected suchthat no primary resonance of the core shaft lies within the runningrange of the engine.

The rearward bearings may be positioned axially level with or rearwardof a leading edge of a lowest pressure turbine blade of the turbine atthe root of the blade.

The rearward bearings may be positioned axially level with or rearwardof a trailing edge of a turbine blade of a third set of turbine bladesfrom the front of the turbine, at the root of the blade, wherein theturbine comprises four sets of turbine blades.

The lowest pressure turbine of the engine has a lowest pressure set ofblades, each blade of the lowest pressure set of blades having a mass,m, a radius at blade mid-height, r, and an angular velocity at cruise,co.

A first blade to bearing ratio of:

$\frac{{the}{minor}{span}(S)}{mr{\omega^{2}\left( {{{for}a{blade}{of}{the}{lowest}{pressure}{set}19c},{19d}} \right)}}$may have a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from4.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and furtheroptionally from 4.5×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

A second blade to bearing ratio of:

$\frac{{the}{minor}{span}(S)}{m \times {r\left( {{{for}a{blade}{of}{the}{lowest}{pressure}{set}19c},{19d}} \right)}}$may have a value in the range from 0.8 to 6.0 kg⁻¹, optionally from 0.8to 5.0 kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0kg⁻¹, optionally from 0.8 to 2.0 kg⁻¹.

The turbine has a length between the leading edge of the forwardmostturbine blade of the turbine and a trailing edge of the rearmost turbineblade of the turbine. A minor span to turbine length ratio may be equalto or less than 1.05. The minor span to turbine length ratio may beequal to or less than 1.00, optionally equal to or less than 0.95. Theminor span to turbine length ratio may be equal to or greater than 0.70,or equal to or greater than 0.75, or equal to or greater than 0.80, orequal to or greater than 0.85. For example the minor span to turbinelength ratio may be in the range from 0.70 to 1.05, optionally from 0.70to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05,optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionallyfrom 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionallyfrom 0.85 to 0.95 in the range from 0.85 to 0.95.

According to a ninth aspect, there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine andthe compressor is the lowest pressure compressor of the engine. Theturbine has a turbine length defined as the distance between the root ofthe most upstream blade of the turbine at its leading edge and the rootof the most downstream blade of the turbine at its trailing edge. Theengine further comprises a fan located upstream of the engine core, thefan comprising a plurality of fan blades; and a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support thecore shaft, the three bearings comprising a forward bearing and tworearward bearings, with a minor span (S) defined as the distance betweenthe two rearward bearings. A minor span to turbine length ratio of:

$\frac{{minor}{span}}{{turbine}{length}}$is equal to or less than 1.05.

The inventor appreciated that keeping the minor span to turbine lengthratio within this range, and more generally smaller than that in knownaircraft, may help to reduce or avoid deleterious whirl modes inoperation.

The minor span to turbine length ratio may be equal to or less than1.00, optionally equal to or less than 0.95. The minor span to turbinelength ratio may be equal to or greater than 0.70, or equal to orgreater than 0.75, or equal to or greater than 0.80, or equal to orgreater than 0.85. For example the minor span to turbine length ratiomay be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00,optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionallyfrom 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to0.95 in the range from 0.85 to 0.95.

The inventor appreciated that the increased length of a core shaft whena gas turbine engine is scaled up may lead to a resonance frequency ofthe core shaft lying in or near the engine running range. Simply scalingup a known engine may therefore lead to increased risks ofresonance-induced damage—the longer core shaft may be problematic. Theinventor discovered that arranging the minor span to be approximatelyequal to, and optionally slightly smaller than, the turbine length (ofthe lowest pressure turbine, in embodiments with more than one turbine),and more specifically within the claimed range, may facilitate avoidanceof deleterious whirl modes, and may reduce damage to the engine in use.

The turbine may comprise a total of four sets of turbine blades. In suchembodiments, both of the two rearward bearings may be located downstreamof the trailing edge of a turbine blade of the third set of turbineblades from the front of the turbine, at the root of the blade.

The two rearward bearings may be located downstream of the leading edgeof the lowest pressure (most downstream) turbine blade of the turbine atthe root of the blade.

The length of the core shaft (L) may be in the range from 1800 mm to2900 mm, optionally from 2000 mm to 2900 mm, further optionally from2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The core shaft may have a running speed range with a lower bound of 1500rpm and an upper bound of 6200 rpm. The upper bound on the core shaftrunning speed range may be the upper bound on maximum take-off (MTO)running speed. The lower bound on the running speed of the core shaftmay be the minimum running speed under ground idle conditions.

The core shaft has a length, L, between the forwardmost bearing and therearmost bearing and a distance (the minor span, S) between the tworearward bearings. The bearings may be arranged such that a length ratio(S/L) of the distance between the two rearward bearings to the coreshaft length is equal to or less than 0.14, or equal to or less than0.13, or equal to or less than 0.12. The length ratio, S/L, of thedistance between the two rearward bearings (S) to the core shaft length(L) may be equal to or greater than 0.05, or equal to or greater than0.06, or equal to or greater than 0.07, or equal to or greater than0.08. For example, the length ratio, S/L, of the distance between thetwo rearward bearings (S) to the core shaft length (L) may be in therange from 0.05 to 0.14, optionally in the range from 0.05 to 0.13,optionally in the range from 0.06 to 0.13, and optionally in the rangefrom 0.08 to 0.13.

The forwardmost bearing of the rearward bearings may have a bearingstiffness in the range of 30 kN/mm to 100 kN/mm. A stiffness ratio ofthe bearing stiffness at the forwardmost rearward bearing divided by thedistance between the two rearward bearings may be in the range from 0.08to 0.5 kN/mm², optionally in the range from 0.08 to 0.40 kN/mm²,optionally in the range from 0.08 to 0.30 kN/mm², optionally in therange from 0.08 to 0.20 kN/mm², optionally in the range from 0.09 to0.40 kN/mm², optionally in the range from 0.15 to 0.50 kN/mm²,optionally in the range from 0.15 to 0.40 kN/mm², and further optionallyin the range from 0.15 to 0.30 kN/mm².

The length (L) of the core shaft may be in the range from 1800 mm to2900 mm or 2750 mm. The distance (S) between the two rearward bearingsmay be in the range from 250 mm to 350 mm, optionally from 260 mm to 350mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The gas length of the core shaft may be in the range from 1800 mm to2900 mm or 2750 mm; and the minor span may be in the range from 250 mmto 350 mm, optionally from 260 mm to 350 mm. The running speed of thecore shaft may be in the range from 1500 rpm to 6200 rpm; and/or adiameter of the fan may be in the range from 330 cm to 380 cm and thegear ratio of the gearbox may be in the range from 3.1 to 3.8. Thelength, minor span and/or running speed may be selected such that noprimary resonance of the core shaft lies within the running range of theengine.

The lowest pressure turbine of the engine has a lowest pressure set ofblades, each blade of the lowest pressure set of blades having a mass,m, a radius at blade mid-height, r, and an angular velocity at cruise,co.

A first blade to bearing ratio of:

$\frac{{the}{minor}{span}(S)}{mr{\omega^{2}\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$may have a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from4.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and furtheroptionally from 4.5×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

A second blade to bearing ratio of:

$\frac{{the}{minor}{{span}{}(S)}}{m \times r\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}$may have a value in the range from 0.8 to 6.0 kg⁻¹, optionally from 0.8to 5.0 kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0kg⁻¹, optionally from 0.8 to 2.0 kg⁻¹.

According to a tenth aspect there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine,and comprises turbine blades, and the compressor is the lowest pressurecompressor of the engine. The engine additionally comprises a fanlocated upstream of the engine core, the fan comprising a plurality offan blades; and a gearbox that receives an input from the core shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft. The engine core further comprises threebearings arranged to support the core shaft, the three bearingscomprising two rearward bearings located downstream of the leading edgeof the lowest pressure turbine blades of the turbine at the root of theblades. The two rearward bearings may therefore be described as beinglocated downstream of the leading edge of the last/rearmost turbineblade.

The inventor appreciated that the core shaft generally moves least (inthe radial direction), and is most level (parallel to the engine axis)at the axial position of the bearings—whilst whirl modes and other bendsor displacements may occur between bearings, the bearings serve to limitradial shaft movement. The inventor appreciated that careful control ofshaft length and bearing position may therefore allow whirl modes of theengine to be managed, so reducing the risk of damage to the engine.

The inventor additionally discovered that positioning the bearingsnearer the bigger and larger turbine stages, towards the rear of theturbine, provides improved turbine support as shaft movements relativeto the turbine position may have more of a deleterious effect on theselarger turbine stages.

The gas turbine engine may further comprise a disc arranged to supportthe lowest pressure turbine blades of the turbine. The two rearwardbearings may be located downstream of a centreline of the disc.

The length of the core shaft may be in the range from 1800 to 2900 mm,optionally in the range from 2000 to 2900 mm, optionally in the rangefrom 2300 to 2800 mm, and further optionally in the range from 2400 to2750 mm.

The core shaft may have a running speed range with a lower bound of 1500rpm and an upper bound of 6200 rpm.

The upper bound on the core shaft running speed range may be the upperbound on maximum take-off (MTO) running speed. The lower bound on therunning speed of the core shaft may be the minimum running speed underground idle conditions.

The core shaft has a length, L, between the forwardmost bearing and therearmost bearing, and a distance, S, between the two rearward bearings.The bearings may be arranged such that a length ratio, S/L, of thedistance between the two rearward bearings (S) to the core shaft length(L) may be equal to or less than 0.14, or equal to or less than 0.13, orequal to or less than 0.12. The length ratio, S/L, of the distancebetween the two rearward bearings (S) to the core shaft length (L) maybe equal to or greater than 0.05, or equal to or greater than 0.06, orequal to or greater than 0.07, or equal to or greater than 0.08. Forexample, the length ratio, S/L, of the distance between the two rearwardbearings (S) to the core shaft length (L) may be in the range from 0.05to 0.14, optionally in the range from 0.05 to 0.13, optionally in therange from 0.06 to 0.13, and optionally in the range from 0.08 to 0.13.

The length (L) of the core shaft may be in the range from 1800 mm to2900 mm or 2750 mm. In such embodiments, the distance between the tworearward bearings (S) may be in the range from 250 mm to 350 mm, oroptionally from 260 mm and 350 mm.

According to an eleventh aspect, there is provided a gas turbine enginefor an aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine,and comprises four sets of turbine blades, and the compressor is thelowest pressure compressor of the engine. The engine further comprises afan located upstream of the engine core, the fan comprising a pluralityof fan blades; and a gearbox that receives an input from the core shaftand outputs drive to the fan so as to drive the fan at a lowerrotational speed than the core shaft. The engine core further comprisesthree bearings arranged to support the core shaft, the three bearingscomprising two rearward bearings located downstream of the trailing edgeof a turbine blade of the third set of turbine blades from the front ofthe turbine, at the root of the blade.

One or more of the following features may apply for a gas turbine engineof the tenth and/or eleventh aspect:

The forwardmost bearing of the rearward bearings may have a bearingstiffness in the range of 30 kN/mm to 100 kN/mm.

A stiffness ratio of the stiffness at the forwardmost rearward bearingto the distance between the two rearward bearings (S) may be in therange from 0.08 to 0.5 kN/mm², optionally in the range from 0.08 to 0.40kN/mm², optionally in the range from 0.08 to 0.30 kN/mm², optionally inthe range from 0.08 to 0.20 kN/mm², optionally in the range from 0.09 to0.40 kN/mm², optionally in the range from 0.15 to 0.50 kN/mm²,optionally in the range from 0.15 to 0.40 kN/mm², and further optionallyin the range from 0.15 to 0.30 kN/mm².

The length (L) of the core shaft may be in the range from 1800 mm to2900 mm or 2750 mm. In such embodiments, the distance between the tworearward bearings (S) may be in the range from 250 mm to 350 mm, oroptionally from 260 mm and 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The length of the core shaft (L) may be in the range from 1800 mm to2900 mm, optionally from 2000 mm to 2900 mm, further optionally from2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

The distance between the two rearward bearings (S), which may bereferred to as the minor span, may be in the range from 250 mm to 350mm. In various embodiments, the minor span may be greater than or equalto any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The running speed of the core shaft may be in the range from 1500 rpm to6200 rpm. Additionally or alternatively, the diameter of the fan may bein the range from 330 cm to 380 cm. The gear ratio of the gearbox may bein the range from 3.1 to 3.8.

The length, minor span and/or running speed may be selected such that noprimary resonance of the core shaft lies within the running range of theengine.

The lowest pressure turbine of the engine has a lowest pressure set ofblades. Each blade of the lowest pressure set of blades has a mass, m, aradius at blade mid-height, r, and an angular velocity at cruise, ω. Aminor span (S) is defined, as mentioned above, as the axial distancebetween the two rearward bearings.

A first blade to bearing ratio of:

$\frac{{the}{minor}{{span}{}(S)}}{mr{\omega^{2}\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$may have a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from4.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and furtheroptionally from 4.5×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

Additionally or alternatively, a second blade to bearing ratio of:

$\frac{{the}{minor}{{span}{}(S)}}{m \times r\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}$may have a value in the range from 0.8 to 6.0 kg⁻¹, optionally from 0.8to 5.0 kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0kg⁻¹, optionally from 0.8 to 2.0 kg⁻¹.

The turbine has a length between the leading edge of the forwardmostturbine blade of the turbine and a trailing edge of the rearmost turbineblade of the turbine. A minor span to turbine length ratio (i.e. Sdivided by the turbine length) may be equal to or less than 1.05,optionally equal to or less than 1.00, optionally equal to or less than0.95. The minor span to turbine length ratio may be equal to or greaterthan 0.70, or equal to or greater than 0.75, or equal to or greater than0.80, or equal to or greater than 0.85. For example the minor span toturbine length ratio may be in the range from 0.70 to 1.05, optionallyfrom 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95,optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and furtheroptionally from 0.85 to 0.95.

The turbine may comprise a total of four sets of turbine blades. In analternative embodiment, the turbine may comprise a total of three setsof turbine blades.

According to a twelfth aspect, there is provided a gas turbine enginefor an aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor.The turbine is the lowest pressure turbine of the engine and thecompressor is the lowest pressure compressor of the engine. The coreshaft has a running speed range with a lower bound of 1500 rpm and anupper bound of 6200 rpm. The engine further comprises a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; and a gearbox that receives an input from the core shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the core shaft.

The engine core further comprises three bearings arranged to support thecore shaft, the three bearings comprising a forward bearing and tworearward bearings. The core shaft has a length (L) between the forwardbearing and the rearmost rearward bearing, and a minor span (S) betweenthe rearward bearings. The bearings are arranged such that a lengthratio of the minor span to the core shaft length (S/L) is equal to orless than 0.14.

The length ratio may be equal to or less than 0.13, or equal to or lessthan 0.12. The length ratio may be equal to or greater than 0.05, orequal to or greater than 0.06, or equal to or greater than 0.07, orequal to or greater than 0.08. For example, the length ratio may be inthe range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13,optionally in the range from 0.06 to 0.13, and further optionally in therange from 0.08 to 0.13.

The inventor appreciated that the increased length of a core shaft whena gas turbine engine is scaled up may lead to a resonance frequency ofthe core shaft lying in or near the engine running range. Simply scalingup a known engine may therefore lead to increased risks ofresonance-induced damage—the longer core shaft may be problematic.

However, making the core shaft still longer—extending rearward of afirst rearward bearing and to a second rearward bearing, with thespacing between the first and second rearward bearings being within aset range with respect to the total core shaft length—was found toincrease stiffness of the core shaft and to move the resonant frequencyaway from the engine running range in some embodiments.

The upper bound on the core shaft running speed range may be the upperbound on maximum take-off running speed.

The lower bound on the core shaft running speed may be the minimumrunning speed under ground idle conditions.

The running speed of the core shaft under cruise conditions may be inthe range from 5400 to 5700 rpm, and optionally in the range from 5500to 5600 rpm.

The running speed of the core shaft under maximum take-off (MTO)conditions may be in the range from 5800 to 6200 rpm, and optionally inthe range from 5900 to 6100 rpm.

The length of the core shaft (L) may be in the range from 1800 mm to2900 mm, optionally from 2000 mm to 2900 mm, further optionally from2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal toany of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, theminor span may be smaller than or equal to any of 350 mm, 345 mm, 340mm, or 335 mm.

The forwardmost bearing of the rearward bearings may have a bearingstiffness in the range of 30 kN/mm to 100 kN/mm. A stiffness ratio ofthe stiffness at the forwardmost rearward bearing to the minor span maybe in the range from 0.08 to 0.5 kN/mm², optionally in the range from0.08 to 0.40 kN/mm², optionally in the range from 0.08 to 0.30 kN/mm²,optionally in the range from 0.08 to 0.20 kN/mm², optionally in therange from 0.09 to 0.40 kN/mm², optionally in the range from 0.15 to0.50 kN/mm², optionally in the range from 0.15 to 0.40 kN/mm², andfurther optionally in the range from 0.15 to 0.30 kN/mm².

The length (L) of the core shaft may be in the range from 1800 mm to2900 mm or 2750 mm. The minor span (S) may be in the range from 250 mmto 350 mm, optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

In some embodiments, the length of the core shaft may be in the rangefrom 1800 mm to 2900 mm or 2750 mm. The minor span may be in the rangefrom 250 mm to 350 mm, optionally from 260 mm to 350 mm. The runningspeed of the core shaft may be in the range from 1500 rpm to 6200 rpm;and/or a diameter of the fan may be in the range from 330 cm to 380 cmand the gear ratio of the gearbox may be in the range from 3.1 to 3.8.In such embodiments, the length, minor span and/or running speed may beselected such that no primary resonance of the core shaft (lies withinthe running range of the engine.

The rearward bearings may be positioned axially level with or rearwardof:

-   -   (i) a leading edge of a lowest pressure turbine blade of the        turbine at the root of the blade; and/or    -   (ii) a trailing edge of a turbine blade of a third set of        turbine blades from the front of the turbine, at the root of the        blade, wherein the turbine comprises four sets of turbine        blades.

The inventor appreciated that positioning the rearward bearings nearerthe bigger and larger turbine stages, towards the rear of the turbine,provides improved turbine support as shaft movements relative to theturbine position may have more of a deleterious effect on these largerturbine stages.

The lowest pressure turbine of the engine has a lowest pressure set ofblades, each blade of the lowest pressure set of blades having a mass,m, a radius at blade mid-height, r, and an angular velocity at cruise,co. A minor span (S) is defined as the axial distance between the tworearward bearings, as stated above.

A first blade to bearing ratio of:

$\frac{{the}{minor}{{span}{}(S)}}{mr{\omega^{2}\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$may have a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from4.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and furtheroptionally from 4.5×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

Additionally or alternatively, a second blade to bearing ratio of:

$\frac{{the}{minor}{{span}{}(S)}}{m \times r\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}$may have a value in the range from 0.8 to 6.0 kg⁻¹, optionally from 0.8to 5.0 kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0kg⁻¹, optionally from 0.8 to 2.0 kg⁻¹.

The turbine has a length between the leading edge of the forwardmostturbine blade of the turbine and a trailing edge of the rearmost turbineblade of the turbine. A minor span to turbine length ratio may be in therange from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05,optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95.

According to a thirteenth aspect, there is provided a gas turbine enginefor an aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and wherein the turbine is the lowest pressure turbine of the engine andthe compressor is the lowest pressure compressor of the engine. Theengine further comprises a fan located upstream of the engine core, thefan comprising a plurality of fan blades; and a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support thecore shaft, the three bearings comprising a forward bearing and tworearward bearings, the distance between the two rearward bearings beingdefined as the minor span, S. The forwardmost rearward bearing has abearing stiffness defined by the radial displacement caused by theapplication of a radial force at the axial centrepoint of the bearing. Astiffness ratio of the bearing stiffness at the forwardmost rearwardbearing to the minor span is in the range from 0.08 to 0.5 kN/mm².

The bearing stiffness ratio may be in the range from 0.09 to 0.40kN/mm², optionally in the range from 0.08 to 0.30 kN/mm², optionally inthe range from 0.08 to 0.20 kN/mm², optionally in the range from 0.09 to0.40 kN/mm², optionally in the range from 0.15 to 0.50 kN/mm²,optionally in the range from 0.15 to 0.40 kN/mm², and further optionallyin the range from 0.15 to 0.30 kN/mm².

The inventor appreciated that controlling the bearing stiffness andminor span such that the ratio of the two lies within the specifiedrange may allow or facilitate management of vibrational modes, sopotentially reducing damage to the engine in use caused by whirl modedisplacements of the core shaft.

The core shaft has a length, L, between the forward bearing and therearmost rearward bearing. The bearings may be arranged such that alength ratio (S/L) of the minor span to the core shaft length is equalto or less than 0.14, or equal to or less than 0.13, or equal to or lessthan 0.12. The length ratio S/L may be equal to or greater than 0.05, orequal to or greater than 0.06, or equal to or greater than 0.07, orequal to or greater than 0.08. For example, the length ratio S/L may bein the range from 0.05 to 0.14, optionally in the range from 0.05 to0.13, optionally in the range from 0.06 to 0.13, and optionally in therange from 0.08 to 0.13.

The core shaft may have a running speed range with a lower bound of 1500rpm and an upper bound of 6200 rpm.

The bearing stiffness of the forwardmost rearward bearing may be in therange of 30 kN/mm to 100 kN/mm. Optionally, bearing stiffness of theforwardmost rearward bearing may be at least substantially equal to 50kN/mm.

The gas turbine engine may further comprise a stationary supportingstructure and a first bearing support structure. The forwardmostrearward bearing may be mounted to the stationary supporting structureby the first bearing support structure. The first bearing supportstructure may be attached to the stationary supporting structure at afirst position located axially rearward of the forwardmost rearwardbearing.

In such embodiments, the first bearing support structure may comprise aplurality of connecting members, which may be spaced circumferentiallyaround the engine axis. The connecting members may connect theforwardmost rearward bearing to the stationary support structure.

Additionally or alternatively, in such embodiments, the first bearingsupport structure may comprise an outer race of the forwardmost rearwardbearing.

In embodiments with a first bearing support structure, the engine mayfurther comprise a second bearing support structure. The second bearingsupport structure may be mounted to the stationary supporting structure,optionally at a second position located forward of, and at a largerradial distance from the engine axis than, the first position. Thesecond bearing support structure may be connected to the first bearingsupport structure by a squeeze film damper in the region of theforwardmost rearward bearing.

The length, L, of the core shaft may be in the range from 1800 mm to2900 mm or 2750 mm. The minor span may be in the range from 250 mm to350 mm, optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The length, L, of the core shaft may be in the range from 1800 mm to2900 mm or 2750 mm and the minor span in the range from 250 mm to 350mm, optionally from 260 mm to 350 mm. The running speed of the coreshaft may be in the range from 1500 rpm to 6200 rpm; and/or a diameterof the fan may be in the range from 330 cm to 380 cm and the gear ratioof the gearbox in the range from 3.1 to 3.8. The length, minor spanand/or running speed may be selected such that no primary resonance ofthe core shaft lies within the running range of the engine.

The rearward bearings may be positioned axially level with or rearwardof a leading edge of a lowest pressure turbine blade of the turbine atthe root of the blade.

The rearward bearings may be positioned axially level with or rearwardof a trailing edge of a turbine blade of a third set of turbine bladesfrom the front of the turbine, at the root of the blade. In suchembodiments, the turbine may comprise four sets of turbine blades.

The lowest pressure turbine of the engine may have a lowest pressure setof blades, each blade of the lowest pressure set of blades having amass, m, a radius at blade mid-height, r, and an angular velocity atcruise, co. The minor span (S) is defined as the axial distance betweenthe two rearward bearings, as described elsewhere.

A first blade to bearing ratio of:

$\frac{{the}{minor}{{span}{}(S)}}{mr{\omega^{2}\left( {{for}a{blade}{of}{the}{lowest}{pressure}{set}} \right)}}$may have a value in the range from 2.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from4.0×10⁻⁶ to 7.5×10⁻⁶ kg⁻¹·rad⁻²·s², optionally from 5.0×10⁻⁶ to 7.5×10⁻⁶kg⁻¹·rad⁻²·s², optionally from 2.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s²,optionally from 3.0×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s², and furtheroptionally from 4.5×10⁻⁶ to 6.5×10⁻⁶ kg⁻¹·rad⁻²·s².

A second blade to bearing ratio of:

$\frac{{the}{minor}{{span}{}(S)}}{m \times r\left( {{{for}a{blade}{of}{the}{lowest}{pressure}{set}19c},{19d}} \right)}$may have a value in the range from 0.8 to 6.0 kg⁻¹, optionally from 0.8to 5.0 kg⁻¹, optionally from 0.8 to 4.0 kg⁻¹, optionally from 0.8 to 3.0kg⁻¹, optionally from 0.8 to 2.0 kg⁻¹.

The turbine has a length between the leading edge of the forwardmostturbine blade of the turbine and a trailing edge of the rearmost turbineblade of the turbine. A minor span to turbine length ratio (minor spandivided by turbine length) may be equal to or less than 1.05, optionallyequal to or less than 1.00, optionally equal to or less than 0.95. Theminor span to turbine length ratio may be equal to or greater than 0.70,or equal to or greater than 0.75, or equal to or greater than 0.80, orequal to or greater than 0.85. For example the minor span to turbinelength ratio may be in the range from 0.70 to 1.05, optionally from 0.70to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05,optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionallyfrom 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionallyfrom 0.85 to 0.95.

In any of the aspects described above, one or more of the followingfeatures may be present:

The turbine may be a first turbine, the compressor may be a firstcompressor, and the core shaft may be a first core shaft. The enginecore may further comprise a second turbine, a second compressor, and aninterconnecting shaft connecting the second turbine to the secondcompressor. The second turbine, second compressor, and second core shaftmay be arranged to rotate at a higher rotational speed than the firstcore shaft.

The engine may further comprise a tail bearing housing located rearwardof the turbine. The tail bearing housing may comprise two bearing discs;each bearing disc may be arranged to support one of the two rearwardbearings. In alternative embodiments, the tail bearing housing maycomprise a single bearing disc, the bearing disc being arranged tosupport one of the two rearward bearings (for example the rearmostbearing).

In embodiments with one or more bearing discs, one or more of thebearing discs may be oriented at least substantially perpendicular tothe engine axis (i.e. at least substantially in a radial plane throughthe engine)

In the various aspects and embodiments described herein, the enginerunning range may be defined as the range of rotation rates of the coreshaft during standard operation of the engine (e.g. during ground idle,take-off, climb and cruise), and may be measured in rotations per minute(rpm). In this context “standard operation of the engine” may excludetransient periods on start-up and shut down, e.g. as core shaft rotationrate increases from zero to the ground idle rotation rate. The enginerunning range includes the ground idle speed, cruise speed, and maximumtake-off (MTO) speed. The core shaft rotation rate may be greater thanor equal to the ground idle rotation rate throughout standard operationof the engine. The core shaft rotation rate during standard operation ofthe engine may also be referred to as the core shaft running speed.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. Thefan tip loading may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20.

The bypass ratio may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 12 to 16, 13 to 15, or 13 to14. The bypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

As used herein, a maximum take-off (MTO) condition has the conventionalmeaning. Maximum take-off conditions may be defined as operating theengine at International Standard Atmosphere (ISA) sea level pressure andtemperature conditions+15° C. at maximum take-off thrust at end ofrunway, which is typically defined at an aircraft speed of around 0.25Mn (i.e. a Mach number of 0.25), or between around 0.24 and 0.27 Mn.Maximum take-off conditions for the engine may therefore be defined asoperating the engine at a maximum take-off thrust for the engine at ISAsea level pressure and temperature+15° C. at an aircraft speed of 0.25Mn.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number, Mn) at mid-cruise of anaircraft to which it is designed to be attached, taking into account thenumber of engines provided to that aircraft. For example where an engineis designed to be attached to an aircraft that has two engines of thesame type, at cruise conditions the engine provides half of the totalthrust that would be required for steady state operation of thataircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a sectional side view of a gas turbine engine illustratingthree bearings on the core shaft;

FIG. 5 is a schematic side view of a gas turbine engine illustrating themajor and minor spans between the bearings;

FIG. 6 illustrates bending modes of the core shaft for two differentbearing configurations;

FIG. 7 is a schematic side view of a gas turbine engine illustrating analternative bearing arrangement to that shown in FIG. 5 ;

FIG. 8 is a schematic view of the mounting of two bearings on a tailbearing housing;

FIG. 9 is a schematic view of the mounting of two bearings on a tailbearing housing of a different arrangement from that shown in FIG. 8 ;

FIG. 10 illustrates a method of operating a gas turbine engine;

FIG. 11 illustrates the first four resonance frequencies of a beam;

FIG. 12 is a close-up view of the bearing housing shown in FIG. 5 ;

FIG. 13 is a perspective view of a first bearing support structure;

FIG. 14 is a cross-sectional view of the first bearing support structureshown in FIG. 13 in position within a bearing housing;

FIGS. 15A and 15B illustrate bearing stiffness determination, inparticular showing the application of a radial force and the resultantdisplacement;

FIG. 16 illustrates a method of obtaining a gas turbine engine asdescribed herein;

FIG. 17 is a graph of displacement against load, illustrating an elasticregion within which stiffnesses of components may be determined;

FIG. 18 illustrates a method of designing a gas turbine engine asdescribed herein;

FIG. 19 is a schematic view of the mounting of two bearings shown inFIG. 9 , with bearing disc angles marked; and

FIG. 20 is a schematic view of a different mounting of two bearings fromthat shown in FIG. 9 , with corresponding bearing disc angles marked.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a core shaft26 and an epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2 . The low pressure turbine 19 (see FIG. 1 ) drives the coreshaft 26, which is coupled to a sun wheel, or sun gear, 28 of theepicyclic gear arrangement 30. Radially outwardly of the sun gear 28 andintermeshing therewith is a plurality of planet gears 32 that arecoupled together by a planet carrier 34. The planet carrier 34constrains the planet gears 32 to precess around the sun gear 28 insynchronicity whilst enabling each planet gear 32 to rotate about itsown axis. The planet carrier 34 is coupled via linkages 36 to the fan 23in order to drive its rotation about the engine axis 9. Radiallyoutwardly of the planet gears 32 and intermeshing therewith is anannulus or ring gear 38 that is coupled, via linkages 40, to a fixedstructure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft with the lowest rotational speedin the engine (i.e. not including the gearbox output shaft that drivesthe fan 23). In some literature, the “low pressure turbine” and “lowpressure compressor” referred to herein may alternatively be known asthe “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3 . Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3 . There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft (core shaft 26), the output shaft and the fixedstructure 24) may have any desired degree of stiffness or flexibility.By way of further example, any suitable arrangement of the bearingsbetween rotating and stationary parts of the engine (for example betweenthe input and output shafts from the gearbox and the fixed structures,such as the gearbox casing) may be used, and the disclosure is notlimited to the exemplary arrangement of FIG. 2 . For example, where thegearbox 30 has a star arrangement (described above), the skilled personwould readily understand that the arrangement of output and supportlinkages and bearing locations would typically be different to thatshown by way of example in FIG. 2 .

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example. In somearrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1 ), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

In the embodiments being described, the engine 10 is a geared gasturbine engine, having a gearbox 30.

The fan 23 is attached to and driven by the low pressure turbine 19 viathe core shaft 26 and an epicyclic gearbox 30. The core shaft 26 issupported by three bearings 26 a, 26 b, 26 c, the three bearingscomprising a forward bearing 26 a and two rearward bearings 26 b, 26 c.In alternative embodiments, more bearings 26 a-c may be provided. Theforward bearing 26 a may be referred to as the first bearing, theforwardmost 26 b of the two rearward bearings 26 b, 26 c as the secondbearing, and the rearmost 26 c of the two rearward bearings 26 b, 26 cas the third bearing.

The forward bearing 26 a is the location bearing 26 a for the core shaft26; i.e. it is the one bearing on the core shaft arranged to limit axialmovement of the core shaft 26 as well as radial movement, so locatingthe core shaft axially. The forward bearing 26 a is mounted on the fixedstructure so as to axially locate the core shaft 26. The skilled personwould appreciate that having multiple location bearings on a singleshaft may cause the shaft to bend deleteriously, or otherwise deform, onexpansion in use, and that use of a single location bearing per shaft istherefore generally favoured.

In the embodiment being described, the forward bearing 26 a is the coreshaft bearing 26 a nearest the front of the engine 10. In alternativeembodiments, there may be one or more other bearings, e.g. rollerbearings, on the core shaft 26 forward of the forward bearing 26 a,however any such bearings are not location bearings—i.e. whilst they mayassist in radially locating the core shaft 26, the core shaft can moveaxially relative to these bearings.

In the embodiments being described, the forward bearing 26 a is the coreshaft bearing 26 a nearest, and rearward of, the gearbox 30. The forwardbearing 26 a is axially level with, or near, the exit from thecompressor 14 in the embodiment being described. In alternativeembodiments, the forward bearing 26 a may be located forward of thegearbox 30, however the proximity of the forward bearing 26 a to aroller bearing of the fan 23 may increase complexity.

The forward bearing 26 a is mounted on the fixed structure.

The rearward bearings 26 b, 26 c are the next two core shaft bearings,rearward of the forward bearing 26 a. In the embodiment being described,the rearward bearings 26 b, 26 c are the core shaft bearings nearest therear of the engine 10. In alternative embodiments, an additional bearingmay be located rearward of these bearings.

In the embodiment shown in FIG. 4 , both rearward bearings 26 b, 26 care mounted on the tail bearing housing 29. The tail bearing housing 29is a structure arranged to be non-rotating with respect to the fixedstructure, and to support bearings 26 b, 26 c of the core shaft 26. Thetail bearing housing 29 comprises two bearing discs 29 a, 29 b. Eachdisc 29 a, 29 b is arranged to support one of the two rearward bearings26 b, 26 c.

In the embodiment shown in FIG. 4 , the rearmost bearing 26 c of the tworearward bearings 26 b, 26 c is located axially level with, or near, theexit from the low pressure turbine 19. More specifically, the rearmostbearing 26 c is at least substantially axially level with arearmost/lowest pressure rotor of the low pressure turbine 19 in theembodiment being described.

In various alternative embodiments, such as that shown in FIGS. 5 to 7 ,one or both of the two rearward bearings 26 b, 26 c may be providedaxially downstream of the leading edge of the lowest pressure set ofblades 19 d of the low pressure turbine 19, at the root of the blade.

In embodiments such as that shown in FIGS. 1 and 4 , the low pressureturbine 19 may have three stages; i.e. three sets of rotor blades 19 a,19 b, 19 c. Each set of rotor blades has a corresponding axial positionalong the engine axis 9, and is offset from the axial position of theother sets. The most upstream, or axially forward-most, set of blades isthe highest pressure set of blades of the low pressure turbine 19, andmay be referred to as the first set of blades. The most downstream, oraxially rear-most, set of blades is the lowest pressure set of blades ofthe low pressure turbine 19, and may be referred to as the last, or inthese embodiments third, set of blades. The middle set of blades 19 b ofthe three sets may be referred to as the second set of blades.

In embodiments such as that shown in FIGS. 5 and 6 , the low pressureturbine 19 may have four stages; i.e. four sets of rotor blades 19 a, 19b, 19 c, 19 d. The most upstream, or axially forward-most, set of bladesis the highest pressure set of blades of the low pressure turbine 19,and may be referred to as the first set of blades 19 a. The mostdownstream, or axially rear-most, set of blades 19 d is the lowestpressure set of blades of the low pressure turbine 19, and may bereferred to as the last, or in these embodiments fourth, set of blades.The middle sets of blades 19 b, 19 c of the four sets may be referred toas the second and third sets of blades, respectively.

In alternative embodiments, the turbine 19 may have less than three ormore than four sets of blades, for example having two sets or five sets.

A length of the low pressure turbine 19 may be defined as the distancebetween a leading edge of a blade 19 a of the first set of blades of thelow pressure turbine and a trailing edge of a blade 19 c/19 d of thelast set of blades of the low pressure turbine. A casing of the lowpressure turbine 19 may extend beyond the span between the first(highest pressure) and last (lowest pressure) blades.

In various embodiments with the two rearward bearings 26 b, 26 cprovided axially downstream of the leading edge of the lowest pressureblade 19 c, 19 d of the low pressure turbine 19 at the root of theblade, the length ratio (as described below) and/or the core shaftrunning speed may or may not be within the ranges detailed elsewhereherein. In such embodiments, the two rearward bearings 26 b, 26 c may beprovided axially downstream of the centreline of a disc supporting thelowest pressure turbine blade 19 d of the low pressure turbine 19, asillustrated in FIG. 12 . In such embodiments, the length, L, of the coreshaft may be in the range 1800-2900 mm, optionally 2300-2800 mm, andfurther optionally 2400-2750 mm.

In various embodiments, the low pressure turbine 19 may be a four stageturbine 19, having four sets 19 a-d of rotor blades, for example asdescribed above with respect to FIGS. 5 and 6 . The forwardmost set ofrotor blades 19 a may be defined as the first set and the rearmost 19 das the fourth set. The two rearward bearings 26 b, 26 c may be providedaxially downstream of the trailing edge of a low pressure turbine bladeof the third set 19 c at its root. One or both of the two rearwardbearings 26 b, 26 c may be provided axially upstream of the leading edgeof the lowest pressure turbine blade 19 d (a blade of the fourth set) ofthe low pressure turbine 19 at the root of the blade in suchembodiments. Pressure decreases across the low pressure turbine 19—thefirst set of blades 19 a may therefore be described as the highestpressure set of blades of the low pressure turbine 19, and the fourthset of blades as the lowest pressure blades.

In various such embodiments, the length ratio and/or the core shaftrunning speed may or may not be within the ranges detailed elsewhereherein. In such embodiments, the length, L, of the core shaft may be inthe range 1800-2900 mm, optionally 2300-2800 mm, and further optionally2400-2750 mm.

In the low pressure turbine 19, turbine blades 19 a-d within each setare designed to be identical (within manufacturing tolerances). Bladesmay differ, for example in size and/or shape, between different sets.Each blade within a set 19 d has a mass, m, and a turbine blade radiusat mid-height, r. Generally, the heavier and larger blades are in thesets towards the rear of the turbine. The turbine blades each have amid-height position, at which the mid-height radius is measured. Themid-height position is half way between a radially innermost point and aradially outermost point on the blade leading edge. The turbine bladeradius at mid-height is measured in a radial direction between an axialcentreline 9 of the engine 10 and the mid-height position.

Each blade 19 a-d also has a maximum-rated angular velocity, ω, whichmay also be referred to as a Maximum Take-Off (MTO) speed. The MTO speedmay be the maximum angular velocity at which the shaft is designed tospin. ω may be in between 5000 and 9000 rpm, optionally in the rangefrom 5000 to 8000 rpm or from 5000 to 7000 rpm, and optionally around5500, 6000 (i.e. around 630 radians per second), or 6500 rpm.

A value, Y, may be defined as follows:Y=mrω ²

Y may have units of kg·m·rad²·s⁻². In some embodiments, the value of Yfor the lowest pressure set of blades 19 c (for a three-stage turbine),19 d (for a four-stage turbine) of the low pressure turbine 19 is in therange of from 45000 to 100000 kg·m·rad²·s⁻², optionally in the range offrom 50000 to 100000 kg·m·rad²·s⁻², optionally in the range of from55000 to 100000 kg·m·rad²·s⁻², and further optionally in the range offrom 60000 to 100000 kg·m·rad²·s⁻². In such embodiments, the mass, m, ofeach blade of the last set may be in the range from 0.2 kg to 0.6 kg,and optionally may be around 0.4 kg. The radius, r, of each blade of thelast set at mid-height may be in the range from 400 mm to 600 mm, andoptionally may be around 500 mm (0.5 m).

The value of Y may be thought of as providing a measure of the magnitudeof the centripetal force (F_(c)) acting on the blade when rotating atthe MTO speed:

$F_{c} = {{ma_{c}} = {\frac{mv^{2}}{r} = {m\omega^{2}r}}}$where a_(c) is the centripetal acceleration and v is the linearvelocity, which is equal to ωr.

The skilled person would appreciate that a higher centripetal force(F_(c)) acting on the blade may increase the risk of a blade-off eventfor that blade.

The minor span, S, is the distance between the two rearmost bearings,and may be in the range from 250 mm to 350 mm, optionally in the rangefrom 275 mm to 325 mm, and further optionally may be around 300 mm.

A first blade to bearing ratio may be defined as:

$\frac{{{the}{minor}{span}},S}{\begin{matrix}{{Y{for}{the}{lowest}{pressure}{set}{of}{blades}19c},} \\{19d{of}{the}{low}{pressure}{turbine}{}19}\end{matrix}}$

The first blade to bearing ratio may be in the range of from 2.0×10⁻⁶ to7.5×10⁻⁶ kg⁻¹·rad⁻²·s², and optionally in the range from 3.0×10⁻⁶ to4.5×10⁻⁶ kg⁻¹·rad⁻²·s². The value of this ratio may be lower than thatfor traditional engines as the minor span, S, is smaller compared to thevalue of Y. The skilled person would appreciate that increasing angularvelocity in flight may improve engine efficiency, and that the MTO speedprovides a measure of a maximum rotation speed, and resultantly amaximum force (indicated by Y), available from an engine 10. However,the inventor appreciated that the engine 10 should not be linearlyscaled up with a force increase (Y providing a measure of force), butrather that the minor span length should be increased as little aspossible so as to relatively reduce engine length and weight, soallowing the efficiency gains to be increased by avoiding the additionalweight, and to avoid the development of unwanted whirl modes within theminor span. Whilst conventional wisdom suggests that a larger minor spanis desirable to improve reaction of forces from the low pressure turbine19, the inventor found that the risk of introducing whirl modes, and theintroduction of more length and weight, counterbalanced the forcereaction benefits.

In alternative or additional embodiments, a second blade to bearingratio may be defined as:

$\frac{{{the}{minor}{span}},S}{{mr}\begin{pmatrix}{{{for}{the}{lowest}{pressure}{set}{of}{blades}19c},} \\{19d{of}{the}{low}{pressure}{turbine}{}19}\end{pmatrix}}$

The value of the blade mass, m, multiplied by the blade radius, r, maybe in the range from 180 to 280 kg·mm. The minor span, S, i.e. thedistance between the two rearmost bearings, may be in the range from 250mm to 350 mm. In various embodiments, the minor span may be greater thanor equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In variousembodiments, the minor span may be smaller than or equal to any of 350mm, 345 mm, 340 mm, or 335 mm. The second blade to bearing ratio may bein the range from 0.8 to 6.0 kg⁻¹, and optionally in the range from 0.9to 3.9 kg⁻¹, and further optionally in the range from 1.2 to 2.6 kg⁻¹.In such embodiments, the gear ratio of the gearbox 30 may be greaterthan 3, and optionally in the range from 3.1 to 3.8.

In such embodiments, the running speed of the engine 10/of the coreshaft 26 may be in the range of 5400-5700 rpm (i.e. around 565 to 597radians per second), and optionally of 5500-5600 rpm, at cruise.Additionally or alternatively, the running speed of the core shaft 26may be in the range of 5800-6200 rpm, and optionally of 5900-6100 rpm,at MTO.

In such embodiments, the length, L, of the core shaft 26 may be in therange 1800-2900 mm, optionally 2000-2900 mm, further optionally2300-2800 mm, and further optionally 2400-2750 mm.

In particular, in the embodiment shown in FIGS. 5 and 6 , both rearwardbearings 26 b, 26 c are mounted on the tail bearing housing 29. The tailbearing housing 29 is located axially rearward of the low pressureturbine 19. Two bearing discs 29 a, 29 b extend from the tail bearinghousing 29. Whilst the forwardmost bearing disc 29 a of the bearinghousing 29 of this embodiment is angled forward of the tail bearinghousing 29, extending towards and, in this case, within the housing of,the low pressure turbine 19, the forwardmost bearing 26 b of therearward bearings, which is mounted on the forwardmost bearing disc 29a, is nonetheless located axially rearward of the leading edge of thelowest pressure turbine blade of the low pressure turbine 19, at theroot of the blade. The rearmost bearing 26 c of the rearward bearings,which is mounted on the rearmost bearing disc 29 b, is located axiallyrearward of the low pressure turbine 19.

In the embodiment shown in FIG. 7 , the tail bearing housing 29 has justone bearing disc 29 c extending therefrom. The single bearing disc 29 cis located in a similar position and has a similar shape to theforwardmost bearing disc 29 a of the embodiment shown in FIGS. 5 and 6 ,but this time supports the rearmost bearing 26 c of the rearwardbearings, holding that bearing 26 c axially rearward of the leading edgeof the lowest pressure turbine blade of the low pressure turbine 19. Theforwardmost bearing 26 b of the rearward bearings is mounted on aforward bearing housing 31 instead of on the tail bearing housing 29.The forward bearing housing 31 has a single bearing disc 31 a arrangedto support the forwardmost 26 b of the rearward bearings. Theforwardmost 26 b of the rearward bearings is located forward of, andadjacent to, the low pressure turbine 19 in this embodiment. Theforwardmost 26 b of the rearward bearings is held axially level with apart of the low pressure turbine 19, and more specifically axially levelwith the leading edge of the highest pressure turbine blade of the lowpressure turbine 19, in this embodiment. The core shaft 26 may beshorter overall in such embodiments with only one rearward bearing 26 crearward of the leading edge of the lowest pressure turbine blade of thelow pressure turbine 19, as the core shaft 26 may not extend as farrearward of the low pressure turbine 19 given that the minor span, S, islocated at least partially axially level with the low pressure turbine19, rather than strictly rearward of the low pressure turbine 19 (morespecifically, of the leading edge of the lowest pressure turbine bladesof the low pressure turbine 19).

A length, L, of the core shaft 26 may be defined between the forwardbearing 26 a and the rearmost rearward bearing 26 c, as marked in FIGS.5 and 6 (i.e. between the first and third bearings). This length, L, maybe referred to as a major span of the core shaft 26, or an inter-bearinglength of the core shaft 26. In the embodiment being described, thelength, L, of the core shaft 26 is in the range 1800-2900 mm, moreparticularly in the range 2300-2800 mm, and still more particularly inthe range 2400-2750 mm.

The minor span, S, of the core shaft 26 may be defined as the distancebetween the rearward bearings 26 b, 26 c (i.e. between the second andthird bearings on the core shaft 26).

The minor span, S, is equal to the length, L, of the core shaft 26 minusthe distance, D, between the forward bearing 26 a and the first rearwardbearing 26 b (i.e. minus the distance between the first and secondbearings on the core shaft 26).

A length ratio of the core shaft may be defined as:

$\frac{{minor}{{span}{}(S)}}{{core}{shaft}{{length}{}(L)}}$

In various embodiments, the bearings 26 a-26 c are arranged such thatthe length ratio of the minor span to the core shaft length is in in therange from 0.08 to 0.14, and optionally in the range from 0.08 to 0.13.In various embodiments, the core shaft 26 may have any suitable lengthratio, for example being in the range of from 0.09 to 0.13, or 0.10 to0.12, or for example being on the order of or at least 0.08, 0.09, 0.10,0.11, 0.12, 0.13 or 0.14. The core shaft length ratio may be, forexample, between any two of the values in the previous sentence.

The inventor appreciated that the ratio of the distance between the reartwo bearings 26 b, 26 c to the shaft length, L, may be a significantparameter in controlling the rotordynamic behaviour of the core shaft26.

The value of this length ratio for engines 10 of various embodiments maybe relatively low compared with previous engines.

As the engine 10 gets bigger, the core shaft 26 gets longer—however,increasing the distance, S, between the two rear bearings 26 b-c (theminor span) linearly with the increase in shaft length, L, may result indeleterious effects. Once the minor span, S, reaches a certain length,there may be no benefit in lengthening that distance in terms of effectson the portion of the core shaft 26 located between the first 26 a andsecond 26 b bearings and/or in terms of reaction to the moments of thebearing discs 29 a-b, 31 a. Indeed, if that distance, S, is increasedlinearly with core shaft length, L, the minor span may become longenough to have whirl modes (as described below with respect to FIG. 11 )between the second and third bearings 26 b, within the engine runningrange, so potentially worsening the unwanted movement.

Additionally or alternatively, a relatively large tail bearing housing(TBH) 29 may be used to accommodate the wider minor span, S, betweenbearings 26 b, 26 c, as shown in FIG. 8 . A larger TBH 29 may addunwanted weight and/or bulk to the engine 10. Alternatively oradditionally, the bearing discs 29 a, 29 b may be more widely angled toaccommodate the wider minor span between bearings 26 b, 26 c, as shownin FIG. 9 . The bearing discs 29 a, 29 b being angled further away froma direction perpendicular to the engine axis 9/parallel to a radius maybe a suboptimal solution from a structural point of view—for example themoment of the bearing discs may not be reacted as well as in theembodiment shown in FIG. 8 . Alternatively or additionally, one of thebearings 29 a, 29 b may be located on a different bearing structure 31,instead of the TBH 29.

In various embodiments, the bearing discs 29 a,b, 31 a, be they on theTBH 29 or on a separate bearing structure 31, may be oriented at leastsubstantially perpendicular to the engine axis 9, for example having anangle (0) to a radius of the engine 10 of between 0° and 20°, andoptionally between 0° and 15° (i.e. an angle to the axis 9 of the engineof between 90° and 70°, and optionally between 90° and 75°).

In embodiments such as that shown in FIGS. 9 and 19 , the bearing discs29 a, 29 b each extend inwards towards the engine axis 9 from the restof the TBH 29 at an at least substantially constant angle (the bearingdiscs 29 a, 29 b may for example comprise a solid disc or a series ofcircumferentially spaced struts around the axis extending between innerand outer rings). An angle, Θ₁, Θ₂, can therefore be defined between thedisc 29 a, 29 b and a radial direction. In other embodiments, such asthose shown in FIG. 20 , a bearing disc 29 a, 29 b may comprise two ormore portions extending at different angles (as shown by way of exampleonly with three portions for the forwardmost bearing disc 29 a and twoportions of the rearward bearing disc 29 b in FIG. 20 ). In such cases,the angle selected is that for the portion with the longest radialextent, as marked for the two examples shown in FIG. 20 . In embodimentsin which there is not a single portion with a greater radial extent thanthe other portion(s), an average angle may instead be taken across thetwo or more portions with the greatest radial extent.

In the embodiments being described, the core shaft 26 has a runningspeed range with a lower bound of 1500 rpm (e.g. at ground idle) and anupper bound of 6200 rpm (e.g. at Maximum Take-Off—MTO). In particular,in various embodiments the core shaft running speed at cruise lieswithin the range 5400-5700 rpm, and optionally within the range5500-5600 rpm. At MTO, the core shaft running speed may lie in the range5800-6200 rpm, and optionally within the range 5900-6100 rpm. Therunning speed range at cruise for a particular aircraft is generallywell below a maximum rated rotation rate for the core shaft 26 of thataircraft (the MTO speed). The engine 10 may operate within the MTOrunning speed range for relatively short time durations—e.g. five or tenminutes—during normal operation.

FIG. 6 illustrates the stiffening effect on the core shaft 26 obtainedby the arrangement of bearings 26 a-26 c as described herein, and inparticular of the use of two rearward bearings 26 b, 26 c. Thepositioning of the bearings is selected to control, or aid incontrolling, the whirl modes of the engine 10.

In particular, the boundary condition on the core shaft 26 at thelocation of the second bearing 26 b (the forwardmost rearward bearing)may be changed by the extension of the core shaft 26 beyond the secondbearing 26 b to the third bearing 26 c (the rearmost rearward bearing).This may change the boundary condition at the second bearing 26 b from asimply-supported/pinned boundary condition to a cantilever boundarycondition, affecting the shape of the core shaft 26 when it bends. Thiseffect can be seen most clearly immediately to the left of the secondbearing 26 b as shown in FIG. 6 —in the upper, two bearing,configuration (pinned-pinned boundary conditions), the angle of the coreshaft at the second bearing 26 b is steeper than in the lower, threebearing, configuration (pinned-cantilever boundary conditions), in whichthe shaft 26 is closer to being horizontal on entering the secondbearing 26 b.

In rotordynamics, the critical speed is the theoretical angular velocitythat excites the natural frequency of a rotating object, such as a shaft(like the core shaft 26). Higher frequency whirl modes may also beinduced (for example at double the natural/resonant frequency). Engines10 of the embodiments being described may be susceptible to having whirlmodes (resonance frequencies) within or near the engine running range,which may cause unwanted and potentially deleterious movement of thecore shaft 26, as is illustrated in FIG. 6 . This coincidence of whirlmodes with the engine running range may be a result of the longer coreshaft 26 of the larger engine 10.

Whilst diameter and/or thickness of the core shaft 26 could be changedto increase the stiffness and therefore push the whirl modes out of theengine running range, the resultant increase in size and/or weight,and/or knock-on effects on other components, may remove or reduce thefeasibility of this option. For example, the core shaft diameter may beconstrained by other engine components located radially outward of thecore shaft 26 (for example, the core shaft 26 is radially inward withrespect to the interconnecting shaft 27 connecting the second (highpressure) turbine 17 to the second (high pressure) compressor 15 in theembodiments being described).

In the embodiments being described, the low pressure turbine 19 drives alow pressure compressor 14 directly, and drives a fan 23 indirectly viaa reduction gearbox 30. A higher pressure system 15, 17, comprising ahigh pressure compressor 15 and a high pressure turbine 17, as well asan interconnecting shaft 27 between them, is located between the lowpressure compressor 14 and the low pressure turbine 19. The core shaft26 is therefore longer than the interconnecting shaft 27 as it extendsacross the full length of the higher pressure system 15, 17 and theadditional length of the low pressure system 14, 19.

The relatively long length, L, of the core shaft 26, as compared toprevious engine architectures, reduces the core shaft's naturalfrequency, bringing the natural frequency within the operational speedof the engine 10, as the natural frequency is inversely proportional tothe major span (core shaft length L).

Natural frequencies exist in various modes (which may be termed whirlmodes), as illustrated in FIG. 11 . A primary resonance (mode 1) has atotal of two nodes (non-moving points), one at each constrained end (thebearings 26 a, 26 b on the shaft 26). A secondary resonance (mode 2) hasone additional node in the centre (a total of three nodes). A thirdresonance mode (mode 3) has four nodes, and a fourth resonance mode(mode 4) has five nodes, etc. The nodes are evenly spaced along the spanof the shaft 26 between the constrained ends 26 a, 26 b. The amplitudedecreases with increasing mode number—the maximum amplitude of the mode1 resonance is greater than that of the mode 2 resonance, etc. Theprimary resonance may therefore be the most damaging, as it causes themaximum radial displacement of the shaft 26. Avoiding operating speedslikely to trigger the primary resonance—by moving the primary resonanceout of the operating speed range—may therefore be of particularinterest.

The relationship of natural frequency of a simply-supported (pinned)beam to the length of the beam is given by:

$f_{n} = {\frac{K_{n}}{2\pi}\sqrt{\frac{EIg}{wl^{4}}}}$where:

-   -   n is the mode number (with mode 1 having nodes at each pinned        end only, as illustrated in FIG. 6 , mode 2 having an additional        node at the mid-point between the ends, etc.);    -   f_(n) is the frequency of the n^(th) mode (the resonance        frequency), measured in Hz;    -   K_(n) is a dimensionless factor dependent on the order of the        mode, n, and also on the applicable boundary conditions, and can        be derived for various boundary conditions as detailed in, for        example, “Roark's Formulas for Stress and Strain,” Warren C.        Young and Richard G. Budynas, Seventh Edition.    -   E is the modulus of elasticity of the beam (N/m²);    -   I is the area moment of inertia of the beam (m⁴);    -   g is the acceleration due to gravity (m/s²);    -   l is the length of the beam between the pinned ends (m); and    -   w is the load per unit length on the core shaft (N/m).

The nodal positions (as a function of length of the beam, l) and valuesof K_(n) for the first five modes may be as tabulated below in Table 1.The data in Table 1 are for a beam with both ends pinned (the boundaryconditions), and K_(n) is therefore equal to (nπ)². For example, for thesecond mode, n=2 and K_(n)=(2π)²=39.5 (to three significant figures).

TABLE 1 Mode K_(n) Nodal position/l 1 9.87 0.0 1.00 2 39.5 0.0 0.50 1.003 88.8 0.0 0.33 0.67 1.00 4 158 0.0 0.25 0.50 0.75 1.00 5 247 0.0 0.200.40 0.60 0.80 1.00

The bending stiffness of the core shaft 26 is a function of both thematerial property and the area moment of inertia, I (geometric). Thearea moment of inertia for a tubular structure, such as the core shaft26, is:

$I = {\frac{\pi}{4}\left( {r_{2}^{4} - r_{1}^{4}} \right)}$where r₁ and r₂ are the inner and outer radii of the tube, respectively.

In order to increase the frequency, the available options include usinga stiffer material for the shaft 26, decreasing the length, L, of theshaft 26, or increasing the shaft diameter. The diameter and the shaftlength are linked to the frequency by a quadratic function and thereforehave the most influence for a change of a given magnitude. However, thediameter is restricted by a lack of available space within the engine10.

Providing a second rearward bearing 26 c modifies the boundaryconditions to mimic a cantilever, as mentioned above—this change inboundary condition has a stiffening effect on the shaft 26, increasingits natural frequency. The change in boundary condition is reflected ina change in K_(n), as shown in Table 2. In particular, the beam now hasjust one end pinned (simply supported), and the other fixed, soK _(n)=π²(n+¼)²

For the example of n=2, K_(n) is therefore (2.25)²π²=50.0 (to threesignificant figures).

TABLE 2 Mode K_(n) Nodal position/l 1 15.4 0.0 1.000 2 50.0 0.0 0.5571.000 3 104 0.0 0.386 0.692 1.000 4 178 0.0 0.295 0.529 0.765 1.000 5272 0.0 0.239 0.428 0.619 0.810 1.000

Assuming all parameters except K_(n) to be constants, it can be seenthat a simple change to the boundary condition could increase the firstEigen frequency by over 50%.

The relative positioning of the core shaft bearings 26 a-c may thereforebe used to tune the stiffening effect to control the resonantfrequencies.

A minor span to turbine length ratio may be defined as:

$\frac{{{minor}{span}},S}{{turbine}{length}}$

In various embodiments, the minor span to turbine length ratio may beequal to or less than 1.05, optionally in the range from 0.85 to 1.05,and further optionally from 0.85 to 0.95.

In such embodiments, the length ratio and/or the core shaft runningspeed may or may not be within the ranges detailed elsewhere herein.Similarly, the rearward two bearings 26 b, 26 c may or may not bepositioned rearward of various blades of the turbine 19 as detailedelsewhere herein. The skilled person will appreciate that variousdifferent approaches for controlling or modifying vibrational modes aredisclosed herein, and that these approaches may be used individually orin any appropriate combination.

In various additional or alternative embodiments, the engine 10 may havea core shaft 26 with a length (L) between the forwardmost and rearmostbearings 26 a, 26 c in the range from 1800 to 2900 mm or 2750 mm, and anaxial separation (minor span, S) between the two rear bearings 26 b, 26c in the range from 250 mm to 350 mm, such that there is no naturalfrequency of the core shaft 26 with only two nodes (one at each end ofthe span between the first and second bearings 26 a, 26 b—a mode 1frequency or primary resonance as shown in Tables 1 and 2 above) withinthe running range of the engine 10.

In such embodiments:

-   -   the core shaft 26 may have a running speed range with a lower        bound of 1500 rpm and an upper bound of 6200 rpm; and/or    -   the fan 23 may have a fan diameter in the range from 330 cm to        380 cm (130-150″) and the gearbox 30 may have a gear ratio        between 3.1 and 3.8.

The core shaft length, L, and the distance, S, between the two rearbearings 26 b, 26 c, and the running speed may therefore be selected sothat the core shaft 26 does not have a primary resonance mode betweenthe first 26 a and second 26 b bearings within the engine running range.A fan diameter of the fan 23 of the engine 10 may be selected to besuitable for a desired running speed—fan diameter may therefore beselected such that an appropriate running speed for the fan does notcause a primary resonance mode of the core shaft 26 between the first 26a and second 26 b bearings within the engine running range.

The skilled person would appreciate that a mode 1 resonance (a primaryresonance) has only two nodes, whereas a mode 2 frequency has anadditional, central, node and a smaller maximum amplitude, A.

When an engine 10 is being designed 3000, prior to manufacture, thedesigner has more scope for adjusting engine parameters. A design method3000 may comprise selecting 3002 positions for the forward bearing 26 aand the forwardmost bearing 26 b of the rearward bearings 26 b, 26 c—forexample selecting the location of each based on the spacing therebetween(e.g. from 1450 to 2500 mm), and on locations of other enginecomponents. The method 3000 may further comprise increasing (ordecreasing) 3004 the length of the core shaft 26 rearward of theforwardmost bearing 26 b of the rearward bearings 26 b, 26 c so as toprovide a suitable shaft length for the minor span (i.e. the distancebetween the two rearward bearings 26 b, 26 c) to be in a suitable range(e.g. from 250 mm to 350 mm) such that there is no primary resonance ofthe core shaft 26 within the running speed range of the core shaft 26.The rearwardmost bearing 26 c of the two rearward bearings (the thirdbearing, in the embodiment being described) may be located at therearward end of the core shaft 26, or adjacent the end of the core shaft26.

In various such embodiments, the core shaft 26/engine 10 may have arunning speed range at cruise of 5400-5700 rpm, and optionally 5500-5600rpm.

In various such embodiments, the core shaft 26/engine 10 may have arunning speed range at MTO of 5800-6200 rpm, and optionally 5900-6100rpm.

In various such embodiments, the core shaft 26 may have a length, L, of1800-2900 mm, optionally 2300-2800 mm, and further optionally 2400-2750mm.

In various embodiments, the bearing stiffness at the forwardmost 26 b ofthe two rearward bearings 26 b, 26 c is controlled. Control of thestiffness may allow or facilitate management of vibrational modes.

The forward of the two rear bearings 26 b (which may also be referred toas the second bearing) has a bearing stiffness in the range of 30 kN/mmto 100 kN/mm, and optionally around 50 kN/mm, through a spring bar (asdescribed below).

The stiffness of the second bearing 26 b is determined, in part, by theengine condition. The lower the excitation at this bearing 26 b, thelower the stiffness offered, but higher the excitation, the higher thestiffness and also the higher the damping. Bearing stiffness is avariable parameter depending on the engine condition, and may, forexample, vary within the range listed above during normal operation ofan aircraft.

The forwardmost rearward bearing 26 b comprises an outer ring 51encircling the core shaft 26; the ring 51 may be referred to as a race,and may contain ball bearings 52 or the likes, as well as oil arrangedto lubricate the race in use. An inner ring 53 (or race), radiallyinward of the outer ring 51, may server to retain the ball bearings 52within the channel formed between the races 51, 53.

The outer race 51 of this rearward bearing 26 b is mounted on thestationary support structure 24 of the engine by means of one or morebearing support structures 50, 55. In the embodiment being described,two bearing support structures 50, 55 are present. The bearing supportstructures 50, 55 may together form a bearing disc 29 a, as shownschematically by comparison of FIGS. 12 and 14 . The bearing supportstructures 50, 55 of the embodiment being described are each connectedto the same tail bearing housing 29. In the embodiment being described,the tail bearing housing 29 forms a part of the stationary supportstructure of the engine 10. The disc 29 a (comprising component supportstructures 50 and 55) extends from the tail bearing housing 29 to thebearing 26 b. In the embodiment being described, the first bearingsupport structure 50 is mounted on the stationary support structure 24of the engine 10 at a first position 58 a. The first position 24 a isaxially rearward of the bearing 26 b in the embodiment being described.The first bearing support structure 50 provides the outer race 51 in theembodiment being described; in particular, the outer race 51 is formedintegrally with the first bearing support structure 50 as shown in FIG.13 . In alternative embodiments, the outer race 51 may be formedseparately from, and mounted on, the first bearing support structure 50.

In the embodiment being described, the first bearing support structure50 comprises a plurality of connecting members 57, which are spacedcircumferentially around the engine axis 9, connecting the outer race 51to a stationary support structure 58. The connecting members 57 extendaxially along a portion of the engine 10.

For example, there may be twenty or forty, evenly-spaced, connectingmembers 57 in some embodiments—numbers and/or spacings may vary in otherembodiments. The one or more connecting members may extend between theouter race of the bearing 26 b, 26 c and a bearing housing 29, 31. Thebearing housing 29, 31 may be rigidly connected to the stationarysupport structure 58, so effectively becoming a part of the supportstructure 58.

The or each connecting member 57 may comprise a metal bar or strut, andmay be referred to as a spring bar. The or each connecting member 57 maybe arranged to provide some flexibility to the bearing 26 b, 26 c, soallowing some radial and/or axial movement in use (e.g. due toexpansion) to be accommodated. The flexibility of the first bearingsupport structure 50 may therefore be referred to as a spring barstiffness.

In the embodiment being described, the second bearing support structure55 is mounted on the stationary support structure 58 of the engine 10 ata second position 58 b. In the embodiment being described, the first 58a and second 58 b positions on the stationary support structure 58 areeach located on a tail bearing housing 29 portion of the stationarystructure 58.

The second position 58 b is axially rearward of the (forwardmost rear)bearing 26 b, but axially forward of the first position 58 a on thestationary support structure 58, in the embodiment being described, andis radially outward from both the bearing 26 b and the first position 58a. In this embodiment, the second bearing support structure 55 isconnected between a radially outer surface of the first bearing supportstructure 50 and the stationary support structure 58, as shown in FIG.14 —in other embodiments, the connection may be different. In theembodiment being described, the second bearing support structure 55 hasa relatively high stiffness (as compared to the first bearing supportstructure 50), and therefore has a negligible contribution to theflexibility of the bearing 26 b. It may be thought of as effectivelyrigid.

A squeeze film damper 56 is provided between the first bearing supportstructure 50 and the second bearing support structure 55, in the regionof the outer race 51. In the embodiment being described, a channel 56 abordered by raised lips is provided around the outer circumference ofthe first bearing support structure 50 to accommodate the squeeze filmdamper 56 and to locate O-ring seals (not shown; axially spaced to lieone at each end of the channel 56 a). In alternative embodiments, nosuch channel and lips may be provided, and/or a channel and lips may beprovided on the second bearing support structure 55 instead or as well,and the squeeze film damper 56 may be (in some cases, entirely)contained by O-ring seals or the likes. This film is in addition to thehydrodynamic oil layer between the bearing 52 and the core shaft 26(i.e. the ball bearings 52, or the likes, are often lubricated by alayer of oil—the squeeze film damper 56 is a layer separate from theball bearings 52).

The squeeze film damper 56 comprises a film layer, usually a layer ofoil or another lubricant, between the bearing 26 b and the housing 30.The squeeze film damper 56 is arranged to soften the bearing support toincrease damping effectiveness. The stiffness of the squeeze film damper56 generally depends on temperature and on shaft rotation rate. Theskilled person would appreciate that any eccentricity in the rotation ofthe bearing 26 b may be damped by the squeeze film damper 56. Thesqueeze film damper 56 may provide some structural isolation of thesupporting structure 58 from the core shaft 26, may reduce theamplitudes of rotor responses to imbalance, and may assist insuppressing rotordynamic instability. This damping may be of particularutility if a blade set has an uneven mass distribution (e.g. due todamage in use, or to a manufacturing defect), or if an unwanted eventduring manufacturing, maintenance or operation moves a blade set out ofradial alignment, such that it is slightly tilted with respect to theengine axis 9.

The inventor appreciated that overall bearing stiffness can be thoughtof as having three parts, which may be considered as springs, forexample arranged in parallel or a combination of series and parallel(depending on component arrangements), namely:

-   -   the stiffness of a layer of oil in the bearing (the squeeze film        damper fluid stiffness);    -   the bearing support 50 stiffness (generally most influenced by        the connecting member 57 stiffness—this may be referred to as a        spring bar stiffness); and    -   the stiffness of the stationary supporting structure 29, 58 on        which the bearing is mounted.

The stationary supporting structure 58 has a much higher stiffness thanthe other two contributions to bearing stiffness, so is treated aseffectively rigid. Any flexibility in the stationary supportingstructure 58 may only become apparent in extreme conditions such asblade-off events. The stiffness of the oil layer varies significantlydepending on shaft rotation rate and temperature, but, at cruise speedsand above, the oil stiffness is much higher than the spring barstiffness of the bearing. The stiffness of each bearing is thereforeconsidered in terms of its spring bar stiffness.

The spring bar stiffness of a bearing 26 b is defined as a radialstiffness—i.e. a linear deflection, δ, along a radius of the engine 10is measured, the deflection being caused by a force, F. This isillustrated in FIGS. 15A and 15B. The diagonal hatching illustrates thatthe stationary structure 58 is deemed to be rigid/unmoving.

FIG. 15A illustrates the bearing support structure 50, 56, 55 (with thecore shaft 26, inner race 53 and ball bearings 52 excluded for clarity),with a portion of a radius, r, of the engine 10 marked on with a dottedline. The marked radial line is located at the axial centrepoint of thebearing race 51. A force, F, is shown applied along the radius, r, in aradially outward direction (i.e. away from the axial centreline). FIG.15B shows the initial position of the first bearing support structure 50(before application of the force, F) in a broken line, and a finalposition of the first bearing support structure 50 (during applicationof the force, F) in a solid line. The skilled person would appreciatethat the deflection shown is much larger than would be expected innormal operation, and is provided for ease of understanding only.Further, the first bearing support structure 50 should return to itsinitial position after removal of the force, F, during normal operation.A displacement, or deflection, δ, is then measured along the radius, r,at the axial centrepoint of the bearing race 51. Two black dotsillustrate the position of the inner surface of the first bearingsupport structure 50 before and during application of the force, F. Thedisplacement, δ, is the distance between those points. The inner surfaceof the first bearing support structure 50 is chosen for ease ofdemonstration only—the skilled person would appreciate that anotherpoint—such as the radially outer surface, or a radial centre point ofthe first bearing support structure 50, or the likes—could be choseninstead. The displacement reflects the combination of compression of thesqueeze film damper 56, bending of the first bearing support structure50 (and in particular of the spring bars 57), and any bending of thesecond bearing support structure 55. The bearing stiffness is thereforea measure of the radial displacement caused by the application of aradial force at the axial centrepoint of the bearing 26 b.

In the embodiment being described, a force, F, of 50 kN causes adeflection, δ, of 1 mm when applied to the forwardmost bearing 26 b ofthe two rear bearings 26 b, 26 c, so the stiffness of that bearing 26 b,as defined herein, is 50 kN/mm. Bearing stiffnesses may vary between thetwo rear bearings 26 b, 26 c, and between embodiments.

The forwardmost bearing of the rearward bearings 26 b therefore has abearing stiffness in the range of 30 kN/mm to 100 kN/mm in theembodiment being described.

A stiffness ratio at the forwardmost rearward bearing 26 b (i.e. thesecond bearing along the core shaft, in the embodiment being described)may be defined as:

$\frac{\begin{matrix}{{the}{bearing}{stiffness}{at}{the}} \\{{forwardmost}{rearward}{}{{bearing}{}\left( {26b} \right)}}\end{matrix}}{{the}{minor}{{span}{}(S)}}$

In various embodiments, this stiffness ratio may be in the range from0.08 to 0.5 kN/mm², and optionally in the range from 0.09 to 0.40kN/mm². Optionally the stiffness ratio may be at least substantiallyequal to 0.25, 0.30, or 0.35 kN/mm². FIG. 10 illustrates a method 1000which may be performed, the method 1000 comprising starting 1002 anengine 10 of an aircraft and reaching cruise conditions, and operating1004 the aircraft under cruise conditions.

The engine 10 may have a running speed range at cruise of 5400-5700 rpm,and optionally of 5500-5600 rpm.

The engine 10 may be operated such that there is no primary resonancemode between the first 26 a and second 26 b bearings under cruiseconditions.

The engine 10 may be operated such that there is no primary resonancemode between the first 26 a and second 26 b bearings anywhere within theengine running range (including both MTO and cruise).

The lengths defined herein, unless otherwise stated, are for thecorresponding component(s) when the engine is off (i.e. at zero speed/onthe bench, at room temperature). These values generally do not varysignificantly over the operating range of the engine (e.g. having only afew mm expansion in shaft length at running temperature, or less); thevalue at cruise conditions of the aircraft to which the engine isattached (those cruise conditions being as defined elsewhere herein) maytherefore be the same as for when the engine is not in use. However,where the length varies over the operating range of the engine, thevalues defined herein are to be understood as being lengths for when theengine is at room temperature and unmoving.

By contrast, as the oil layer stiffness is speed-dependent, andcontributes to bearing stiffness, the bearing stiffness is defined atcruise conditions/with the shaft rotating at a suitable speed forcruise.

FIG. 17 illustrates how the bearing stiffnesses defined herein may bemeasured. FIG. 17 shows a plot of the displacement δ resulting from theapplication of a load L (e.g. a force, moment or torque) to an arbitrarycomponent for which the stiffness is being measured. At levels of loadfrom zero to L_(P) there is a non-linear region in which displacement iscaused by motion of the component (or relative motion of separate partsof the component) as it is loaded, rather than deformation of thecomponent; for example moving within clearance between parts. For abearing 26 b, the amount of displacement possible in this non-linearregion is likely to be very small. At levels of load above L_(Q) theelastic limit of the component has been exceeded and the applied load nolonger causes elastic deformation—plastic deformation or failure of thecomponent may occur instead. Between points P and Q the applied load andresulting displacement have a linear relationship. The stiffnessesdefined herein may be determined by measuring the gradient of the linearregion between points P and Q (with the stiffness being the inverse ofthat gradient). The gradient may be found for as large a region of thelinear region as possible to increase the accuracy of the measurement byproviding a larger displacement to measure. For example, the gradientmay be found by applying a load equal to or just greater than L_(P) andequal to or just less than L_(Q). Values for L_(P) and L_(Q) may beestimated prior to testing based on materials characteristics so as toapply suitable loads.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention claimed is:
 1. A gas turbine engine for an aircraftcomprising: an engine core comprising a turbine, a compressor, and acore shaft connecting the turbine to the compressor, the turbine being alowest pressure turbine of the gas turbine engine and the compressorbeing a lowest pressure compressor of the gas turbine engine; and a fanlocated upstream of the engine core, the fan comprising a plurality offan blades; wherein: the turbine comprises a total of three sets ofturbine blades, the engine core further comprises three bearingsarranged to support the core shaft, the three bearings comprising aforward bearing and two rearward bearings located downstream of aleading edge of a lowest pressure turbine blade of the turbine at a rootof the blade, the fan comprises a fan diameter in the range from 240 cmto 280 cm, the core shaft has a length, L, between the forward bearingand a rearmost rearward bearing of the two rearward bearings and a minorspan, S, between the two rearward bearings, and the three bearings arearranged such that a length ratio, S/L, of the minor span, S, to thelength, L, of the core shaft is equal to or less than 0.14.
 2. The gasturbine engine of claim 1, wherein the length ratio, S/L, is in therange from 0.08 to 0.13.
 3. The gas turbine engine of claim 1, whereinthe length of the core shaft is in the range from 1800 mm to 2900 mm. 4.The gas turbine engine of claim 1, wherein a forwardmost rearwardbearing of the two rearward bearings has a bearing stiffness in therange of 30 kN/mm to 100 kN/mm.
 5. The gas turbine engine of claim 1,further comprising a gearbox that is configured to receive an input fromthe core shaft and output drive to the fan so as to drive the fan at alower rotational speed than the core shaft.
 6. The gas turbine engine ofclaim 5, wherein the gearbox has a gear ratio greater than
 3. 7. The gasturbine engine of claim 6, wherein the gear ratio of the gearbox is inthe range from 3.1 to 3.8.
 8. The gas turbine engine of claim 1, furthercomprising a tail bearing housing located rearward of the turbine andcomprising two bearing discs, each bearing disc being arranged tosupport one of the two rearward bearings.
 9. The gas turbine engine ofclaim 8, wherein the two bearing discs are oriented at leastsubstantially perpendicular to an axis of the gas turbine engine. 10.The gas turbine engine of claim 1, wherein: the turbine is a firstturbine, the compressor is a first compressor, and the core shaft is afirst core shaft; the engine core further comprises a second turbine, asecond compressor, and a second core shaft connecting the second turbineto the second compressor; and the second turbine, second compressor, andsecond core shaft are arranged to rotate at a higher rotational speedthan the first core shaft.
 11. A gas turbine engine for an aircraftcomprising: an engine core comprising a turbine, a compressor, and acore shaft connecting the turbine to the compressor, the turbine being alowest pressure turbine of the gas turbine engine and the compressorbeing a lowest pressure compressor of the gas turbine engine; and a fanlocated upstream of the engine core, the fan comprising a plurality offan blades; wherein: the turbine comprises a total of three sets ofturbine blades, the engine core further comprises three bearingsarranged to support the core shaft, the three bearings comprising aforward bearing and two rearward bearings located downstream of aleading edge of a lowest pressure turbine blade of the turbine at a rootof the blade, the fan comprises a fan diameter in the range from 240 cmto 280 cm, a forwardmost rearward bearing of the two rearward bearingshas a bearing stiffness in the range of 30 kN/mm to 100 kN/mm, the coreshaft has a minor span between the two rearward bearings, and astiffness ratio of the bearing stiffness at the forwardmost rearwardbearing to the minor span is in the range from 0.08 to 0.5 kN/mm².
 12. Agas turbine engine for an aircraft comprising: an engine core comprisinga turbine, a compressor, and a core shaft connecting the turbine to thecompressor, the turbine being a lowest pressure turbine of the gasturbine engine and the compressor being a lowest pressure compressor ofthe gas turbine engine; and a fan located upstream of the engine core,the fan comprising a plurality of fan blades; wherein: the turbinecomprises a total of three sets of turbine blades, the engine corefurther comprises three bearings arranged to support the core shaft, thethree bearings comprising a forward bearing and two rearward bearingslocated downstream of a leading edge of a lowest pressure turbine bladeof the turbine at a root of the blade, the fan comprises a fan diameterin the range from 240 cm to 280 cm, the turbine has a length between aleading edge of a forwardmost turbine blade of the turbine and atrailing edge of a rearmost turbine blade of the turbine, the core shafthas a minor span between the two rearward bearings, and a ratio of theminor span to the length of the turbine is equal to or less than 1.05.13. The gas turbine engine of claim 12, wherein the ratio of the minorspan to the length of the turbine is in the range from 0.70 to 1.05. 14.A gas turbine engine for an aircraft comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor, the turbine being a lowest pressure turbineof the gas turbine engine and the compressor being a lowest pressurecompressor of the gas turbine engine; a fan located upstream of theengine core, the fan comprising a plurality of fan blades; and a gearboxthat is configured to receive an input from the core shaft and outputdrive to the fan so as to drive the fan at a lower rotational speed thanthe core shaft, wherein: the turbine comprises a total of three sets ofturbine blades, the engine core further comprises three bearingsarranged to support the core shaft, the three bearings comprising aforward bearing and two rearward bearings, the core shaft has a minorspan, S, between the rearward bearings, the turbine has a length betweena leading edge of a forwardmost turbine blade of the turbine and atrailing edge of a rearmost turbine blade of the turbine, a minor spanto turbine length ratio is equal to or less than 1.05, and the fancomprises a fan diameter in the range from 240 cm to 280 cm.
 15. The gasturbine engine of claim 14, wherein the minor span to turbine lengthratio is in the range from 0.70 to 1.05.